The present invention relates to a gas turbine engine and, more particularly, to thermal management of the bearing system for a miniature gas turbine engine.
Miniature gas turbine or turbojet engines (typically of 150 lbf thrust and smaller) are often utilized in single usage applications such as reconnaissance drones, cruise missiles, decoy and other weapon applications, including air-launched and ground-launched weapon systems. The use of such an engine greatly extends the range of the weapon in comparison to the more conventional solid fuel rocket engine.
To achieve economically feasible extended range expendable propulsion sources for such applications, it is necessary that the miniature gas turbine engines be manufactured relatively inexpensively yet provide the highest degree of starting and operational reliability when launched from air or ground systems. One component that greatly affects mechanical performance and reliability is the high speed bearings which support the rotating turbine machine. Reliability and efficiency of the bearing system is a priority for a successful expendable turbine engine. Such reliability and efficiency of the bearing system may be compromised through foreign object damage (FOD), inadequate thermal management, or inadequate lubrication distribution.
Current gas turbine bearing systems employ a relatively complex closed circuit lubrication scheme which is relatively expensive to manufacture and difficult to maintain over long term storage typical of single use systems. Other gas turbine bearing systems utilize open air flow-through systems which, although providing satisfactory thermal management, directly subject the rotating components to FOD which may increase the potential of an operational failure.
Accordingly, it is desirable to provide an uncomplicated and inexpensive thermal management and lubrication system for a miniature gas turbine engine which facilitates storage yet assures operational reliability.